The invention relates to a method for repairing aircraft primary structures. More specifically, the invention relates to a method for repairing damage to the coating of an aircraft fuselage, which coating is made of fiber-reinforced composite material.
While in operation, aircraft structures are frequently subjected to impacts with birds, vehicles, tools, airport installations, etc. Such impacts create damaged areas within the structure, which have lower mechanical resilience; these areas become prime sites for the start of defects such as cracks, which are likely to propagate.
More specifically, in the case of a fuselage made of fiber-reinforced composite material, the actual damaged area may be significantly larger than the visible impact area.
When an aircraft is subjected to such damage, it must absolutely be repaired so as to recover all the structural strength and avoid the propagation of defects from the damaged area before it can be put back into service.
It must be possible to carry out such repairs within the shortest possible immobilization time for the plane.
A known repair mode involves covering the damaged area with a generally circular liner or patch, whose area is significantly larger than said damaged area, then fastening it to the portion of the structure that remains sound with any appropriate means, such as rivets, bolts, welding or bonding.
This solution has the advantage of being relatively simple to implement. It does, however, have the disadvantage of keeping the damaged zone, which remains under loading and therefore may initiate the propagation of defects, even though the presence of the liner limits the flow of forces to which said area is subjected. Consequently, such repairs can only be very temporary and need to be monitored very closely until a permanent repair can be effected.
In addition, the liner must be shaped to match the shape of the fuselage in the area under consideration. This shape may be complex and non-involute, such that it requires specific shaping of the liner, which being circular in shape must then be cut-out of a larger-size plate, itself made by any sheet metalworking means.
According to another embodiment described, for example, in international patent application WO2007135318 in the name of the applicant, a polygonal cut-out is made around the impact area so as to eliminate the whole of the damaged area. A liner, also polygonal in shape but with a larger surface area, is then fastened to the part that remains sound so as to close the cut-out. This method avoids the initiation of defects in the damaged area, since this has been eliminated. Nevertheless, it is then necessary to ensure that the cutting-out operation itself does not introduce any defects. In the case of patent application WO2007135318, holes centered on the intersection of the sides of the polygon are drilled at each corner of said parallelepiped before cutting-out to facilitate this and to avoid unfortunate “saw-cuts” in the delicate cut-outs at the corners. The method divulged in this patent application aims therefore to avoid the formation of cutting defects which may give rise to cracks at the corners of the polygonal cut-out. Effectively, it is in these areas that the beginnings of cracks are most likely to occur, subsequent to a lack of precision on the part of the operator assigned to cutting-out.
In the case of a fiber-reinforced composite material fuselage, defects such as delamination can be caused during cutting-out, irrespective of the care taken by the operator. The action of the cutting implement, of whatever kind it may be, can easily break the cohesion of plies located at the edge of the cut-out and thus initiate delamination. Where the material that is cut-out is a fiber-reinforced composite material, the risk of introducing defects at the edge of the cut-out depends essentially on the rate of advance of the tool. At a given cutting speed, too high a rate of advance favors the occurrence of delamination, whereas too slow a rate of advance causes thermal degradation of the matrix (burning or melting). This type of cutting-out operation in fiber-reinforced composite materials must therefore be performed at a controlled rate of advance, within a narrow band of admissible speeds. In the case of repair operations, however, the cutting-out advance movement is generally communicated manually to the machine by the actions of the operator who moves it along the path to be cut. Even if a judicious choice of tool geometry can increase the range of favorable cutting-out conditions, it remains difficult or even impossible to prevent delamination of the surface plies from occurring, except where sophisticated methods of automatic advance are implemented, which control the cutting parameters and, in particular, the rate of advance/cutting speed combination along complex trajectories that match the shape of the fuselage.
There is therefore a requirement for a repair method for an aircraft fuselage made of composite material that can be implemented with adequate safety in conditions compatible with the means of airport maintenance workshops and requiring the shortest possible immobilization time for the plane.